Intelligent multifunctional actuation system for vibration and buffet suppression

ABSTRACT

Circuits and methods for use on a mobile platform including a flight control system, a structure, an aerodynamic surface, and an actuator operatively coupled to the surface to control the surface. The circuit includes a first and a second input, a summing element, and an output. The first input accepts commands from the flight control system while the second input accepts a vibration signal from the structure. The summing element communicates with the inputs and sums the signal and the command. In turn, the summing element controls the actuator with the summed signal and command.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. patent application Ser. No.10/798,687 filed on Mar. 11, 2004. The entire disclosure of the aboveapplication is incorporated herein by reference.

FIELD

This invention relates generally to flight control systems and, moreparticularly, to flight control systems for suppressing aerodynamicallyinduced vibrations.

BACKGROUND

High performance fighter aircraft such as the F-18 and F-22 (availablefrom the Boeing Company of Chicago, Ill.) often experience highfrequency aerodynamically induced vibrations of their wings,stabilators, and vertical tails. These vibrations are cause by buffetingaerodynamic forces and are transmitted into, and through, the aircraftstructure. If uncompensated for, the associated stresses may lead topremature cracking of the structure. The resulting repairs of thesecracks are both expensive and time consuming. In the alternative,aerodynamic modifications to reduce the causative turbulence imposeperformance constraints on the aircraft while structural modificationsto reduce the resulting fatigue stress incur weight and cost penalties.

Additionally, the buffeting of the vehicle transmits noise into, andcauses noise within, the aircraft structure. In turn, the structuretransmits the noise to the cockpit wherein noise control measures, withadditional penalties must be undertaken. Nor are these problems isolatedto high performance aircraft. Automobiles, missiles, and launch andreentry vehicles (for example) also receive buffeting from aerodynamicforces.

Various attempts have been made to use the existing flight controlactuators to compensate for these aero-vibrations. However, the flightcontrol system typically commands the actuator at rates of about 30cycles per second or less. Since the vibrations occur at frequenciessignificantly higher than the commands, such attempts have failed.

Moreover, attempts to modify the flight control system to react quicklyenough to respond to the vibration are impractical for a variety ofreasons. For instance, such modifications require an order of magnitudeincrease in the flight computer speed. Thus, these solutions necessitatean upgrade of the computer. Additionally, modifying the flight controllaws to accommodate the additional functionality necessitate therecertification of the flight control system. These recertifications areexpensive, time consuming, and (as such) highly undesirable.

Furthermore, as composite materials replace aluminum, and otherconventional, structural members (e.g. on the Boeing 7E7 aircraft)vibration control may assume an increasing importance in the design,operation, and maintenance of aircraft. Thus a need exists to reduce oreliminate aerodynamically induced vibrations.

SUMMARY

It is in view of the above problems that the present invention wasdeveloped. The invention provides apparatus and methods for reducingvibrations of mobile platform structures.

In one preferred embodiment, the invention provides a self-containedactuation device that reduces the aerodynamic buffet loads. Accordingly,construction and operation of mobile platforms (e.g. aircraft) inaccordance with the principles of the present invention results in morecost efficient platforms that possess better performance and longerservice lives.

Another preferred embodiment employs the existing flight controlactuators to reduce the aero-buffeting without requiring modification ofthe flight control system. The present embodiment also includes avibration sensor placed on an aircraft structure (e.g. a wing) to sensethe aero-vibration. The sensor is connected to a controller that iscoupled to the actuator. The controller inverts the vibration signal andadds it to the actuator command from the flight control system. Then thecontroller sends the combined signal (the command with the invertedvibration signal superimposed thereon) to the actuator, thereby drivingthe actuator out of phase with the vibration. By thus canceling thevibration the current embodiment reduces cyclical loads and improves thefatigue life of the structure.

In yet another preferred embodiment, a method of reducingaerodynamically induced vibrations is provided. The method includessensing the vibration and inverting a signal representative of the same.The inverted vibration signal is superimposed on a flight control systemcommand for an actuator to drive the actuator out of phase with thevibration.

Further features and advantages of the present invention, as well as thestructure and operation of various embodiments of the present invention,are described in detail below with reference to the accompanyingdrawings.

DRAWINGS

The accompanying drawings, which are incorporated in and form a part ofthe specification, illustrate the embodiments of the present inventionand together with the description, serve to explain the principles ofthe invention. In the drawings:

FIG. 1 illustrates an aircraft constructed in accordance with theprinciples of the present invention;

FIG. 2 is a detailed view of a wing of the aircraft of FIG. 1;

FIG. 3 shows an actuator in accordance with the principles of apreferred embodiment of the present invention;

FIG. 4 is a block diagram of a control system in accordance with theprinciples of another preferred embodiment of the present invention;

FIG. 5 is a schematic diagram of a circuit in accordance with apreferred embodiment of the present invention; and

FIG. 6 is a flowchart of a method in accordance with a preferred form ofthe present invention.

DETAILED DESCRIPTION

Referring to the accompanying drawings in which like reference numbersindicate like elements, FIG. 1 illustrates a pair of aircraft 10constructed in accordance with a preferred embodiment of the presentinvention. The aircraft 10 includes a number of structures and controlsurfaces that are well known in the art and will herein be representedby an exemplary wing 12 and an aileron 14. Additionally, a fairing 18that covers an aileron actuator may be seen on the underside of the wing12.

During flight, the flight computer and the pilot continually sense theflight related conditions (pitch, climb rate, speed, and the like).Depending on the conditions, either the computer, the pilot, or bothissue commands (e.g. electromagnetic signals) to reposition the ailerons14 and other control surfaces. These commands cause the actuator 20 toeither extend or retract to move the aileron 14. As noted previously,the wing 12 rushing through the air causes turbulence in the air thatcauses the wing 12 to vibrate. Additionally, the movement of the controlsurface 14 tends to change the airflow, thus introducing another sourceof turbulence, buffeting, and vibration.

Referring now to FIG. 3, an actuator 20 for the aileron 14 is shown. Theactuator 20 includes an attachment fitting 22 for coupling the actuator20 to the wing 12 and a ram 24 that moves in response to flight controlsystem commands. The ram 24 is operatively coupled to the aileron 14 tochange the aileron position. Additionally, FIG. 3 shows a circuit 26being mechanically coupled to the actuator 20 as well as a vibrationsensor 28 rigidly coupled to the actuator 20. While FIG. 3 shows thevibration sensor 28 separate from the circuit 26, the sensor 28 may beincluded in the circuit 26. It should also be noted that the actuator 20may also include an internal position sensor such as an LVDT (linearvariable differential transformer) to detect the position of the ram 24.

Additionally, the actuator typically includes a conventional power cable30 for receiving power from the aircraft 10 power subsystem. A secondpower cable 32, in accordance with the principles of the presentinvention, branches from the first power cable 30 to supply the circuit26 power. Importantly, branching cable 32 from cable 30 saves cableweight by obviating the need for a dedicated cable run from the aircraft10 power subsystem to power the circuit 26. Another cable 30′ carriesthe flight control system command to the circuit 26, while in previoussystems the cable 30′ was connected directly to the actuator instead ofthe circuit.

Moreover, coupling the vibration sensor 28 to the actuator 20 instead ofthe wing 12 (FIG. 1) eliminates the need for, and associated weight of,a vibration signal cable from the wing to the circuit 26. Though, thevibration sensor 28 may be located on the wing 12 without departing fromthe spirit or scope of the invention. Additionally, the elimination ofthe wing-to-circuit vibration signal cable simplifies the interfaceswith the wing since a cable 31 on the actuator suffices to carry thevibration signal to the circuit. In turn, the simplification reducesairframe integration costs. Similarly, integrating the vibration sensor28 with the circuit 26, rather than mounting the two on the actuator 26separately, simplifies construction and installation of the actuator andreduces costs further still.

Furthermore, coupling the circuit 26 to the actuator 20 eliminates theneed for cables between the circuit and the actuator to carry thecommand signal from the circuit and to carry the position signal fromthe actuator. Likewise, because the flight control system alreadyincludes a cable 30′ to the actuator (to carry the actuator command) thepresent invention requires no modifications to the flight controlsystem. Accordingly, the present invention further reduces integrationcosts and delays.

With reference now to FIG. 4, the block diagram shown therein representsa system in accordance with an exemplary embodiment of the presentinvention. The system 100 includes a summing element 102, an actuator,104, a control surface 106, a structure 108 (e.g. a wing), a vibrationsensor 110, an invertor 112, a filter 114, and an actuator positionsensor 116. Upon arrival from the flight control system a command passesthrough the summing element 102 (assuming for the moment there is nocurrent vibration). The command causes the actuator 104 to move andreposition the control surface 106. Because of aerodynamic turbulenceover the control surface 106 and wing 108, the wing begins to vibrate.In turn, the vibration sensor 110 generates a signal representative ofthe vibration that the inverter 112 inverts.

To complete the feedback loop, the summing element 102 superimposes theinverted vibration signal on the command. The resulting signal (thecommand with the inverted vibration signal superimposed on it) is fed tothe actuator 104 to drive the control surface out of phase with thevibration. Accordingly, the control surface 106 causes a disturbancethat cancels the vibration of the wing 108. Those skilled in the artwill recognize that the commands typically operate up to approximately30 Hz while the vibration occurs at substantially higher frequencies.Similarly, the commands are typically signals having amplitudes well inexcess of the amplitude of the vibration signal (or can be so tailoredwith appropriate choice of system gains). Thus the inverted vibrationsignal appears as a ripple superimposed upon the command signal afterthe two are added.

Since many flight control systems are designed to sense the actualposition of the actuator 104, a feedback signal may also be provided bythe present invention. In particular the vibration sensor 110 and theposition sensor 116 (of the actuator 104) may communicate with thefilter 114 that subtracts the vibration signal from the position signalas sensed by the position sensor. Accordingly, the position signal fedback to the flight control system does not reflect the slight differencebetween the commanded position and the actual position as influenced bythe vibration canceling circuit 126. Thus, by filtering the vibrationsignal from the actuator position signal, the present invention ensuresthat the flight control system operates properly (e.g. does not raise afalse alarm to indicate that the actuator is out of position).

Note should also be made that FIG. 4 depicts a circuit 126. The circuit126 includes the summing element 102, the inverter 112, and perhaps thefilter 114 connected as shown. Hence, the circuit 126 combines thecommand and the inverted vibration signal to cancel the vibration. Thecircuit may also provide the filtered position signal as shown in FIG.4. In one preferred embodiment, the control surface 106 is an aileronand the vibration sensor is placed as close to the wing tip as ispractical. In this manner, the flexibility of the wing and the positionof the sensor 110, well out on the wing ensure that a robust vibrationsignal is generated.

Now with reference to FIG. 5, a simplified schematic of a circuit 226 inaccordance with a preferred embodiment of the present invention isillustrated. The circuit 226 includes a summing amplifier 202, aninverting amplifier 204, a difference amplifier 206, and an internalpower supply 208. Interfaces with the circuit 226 include a flightcontrol system command input 210, a vibration signal input (here, thevibration sensor 212 is shown as being internal so that the interface isnot necessary), a command output 214, an actuator position signal input216, and a position signal output 218. Internally, an inverted vibrationsignal 220 is also shown.

Additionally, a flight control actuator 228 is shown external to thecircuit 226. Schematically FIG. 5 illustrates the actuator 228 as asolenoid to indicate that the actuator contains components that requireelectric power. The required electric power is typically supplied by theaircraft power subsystem 224 via a cable 230. A second power cable 232is shown branching from the cable 230 and providing power to the powerinput 222 for the circuit 226. Thus, the circuit 226 and actuator sharea common power supply. Those skilled in the art will recognize that theinternal power supply 208 is shown in simplified form for clarity.Details such as neutral returns and interconnections to the amplifiers202 to 206 have been likewise omitted.

Additionally, a housing 234 of the circuit 226 is rigidly coupled to theactuator so that the vibration transducer 212 accurately senses thevibration of the actuator. In turn, the actuator is rigidly mounted tothe structure for which vibration reduction is desired (e.g. by theattachment means 22 shown by FIG. 3). Because of the rigid couplingbetween the structure and the actuator, and between the actuator and thesensor 212, the sensor 212 provides a reliable indication of thevibration of the structure.

Thus, in operation, the summing amplifier 202 sums the flight controlcommand 210 and the vibration signal 220 as inverted by the invertingamplifier 204. The summing amplifier 202 outputs the command 214, withthe inverted vibration signal superimposed thereon, to the actuator. Asnoted previously, the actuator causes an aerodynamic disturbance thatcancels the vibration of the structure. Subsequently, if anothertransient disturbance causes the vibration to return, the vibrationcancellation loop (as just described) acts to cancel the new vibration.In the alternative, a non-inverted vibration signal could be subtractedfrom the command to the same general effect of canceling the vibration.

The circuit 226 also includes a position feedback subsystem thatincludes the difference amplifier 206. The amplifier 206 accepts theposition signal 216 from the actuator position sensor and subtracts thesignal from the vibration signal supplied by the sensor 212 therefrom.Accordingly, the amplifier 206 filters the vibration signal from theposition signal 216 and communicates the result 218 to the flightcontrol system.

With reference now to FIG. 6, a flowchart illustrating a method inaccordance with the principals of the present invention is shown. Themethod includes mounting a vibration canceling circuit to either thewing or the aileron actuator as shown at 302 and 304 respectively. Inoperations 306 and 308 a vibration sensor is coupled to either the wingor the actuator. Jumpers are then installed between the circuit and thevibration sensor, the circuit and a position sensor of the actuator, andthe circuit and the actuator power supply (preferably within theenvelope of the actuator). See operation 310. Moreover, the flightcontrol system cable that carries the actuator command is moved from theactuator to the circuit in operation 311.

With the power on, commands for the actuator may then be received whilethe vibration is sensed as at 312 and 314. The two signals are added,subtracted, or superimposed to obtain a command with an invertedvibration signal superimposed upon it at operation 315. The combinedsignal, from operation 315, is used to drive the actuator out of phasewith the vibration. See operation 316. Meanwhile, the position of theactuator may be sensed in 318. Accordingly, the vibration signal may befiltered from the position feedback signal and the result forwarded theflight control system. See for instance operation 320. In this manner,the method may repeat for as long as vibration canceling is desired asindicated at decision 322.

In view of the foregoing, it will be seen that the several advantages ofthe invention are achieved and attained. Notably, the embodimentsdescribed herein require no added cabling, or cabling modificationsoutside of the envelope of the actuator. Moreover, the present inventionrequires no modification to the flight control system, or even theflight computer. Thus, the present invention provides vibrationelimination with a weight and cost savings over previous attempts toreduce aerodynamically induced vibrations.

The embodiments were chosen and described in order to best explain theprinciples of the invention and its practical application to therebyenable others skilled in the art to best utilize the invention invarious embodiments and with various modifications as are suited to theparticular use contemplated. For example, many industrial machinesinclude structures that are moved by an actuator commanded by a controlsystem. If the machine includes a tool (e.g. a drill) on the movablestructure, the tool may induce undesired vibrations in the structure.Thus, the apparatus and methods discussed herein may be adapted to sensethe tool induced vibration of the structure and cancel the same withoutrequiring modification of the machine control system. Accordingly,machines with increased precision and accuracy are also provided by thepresent invention.

As various modifications could be made in the constructions and methodsherein described and illustrated without departing from the scope of theinvention, it is intended that all matter contained in the foregoingdescription or shown in the accompanying drawings shall be interpretedas illustrative rather than limiting. Thus, the breadth and scope of thepresent invention should not be limited by any of the above-describedexemplary embodiments, but should be defined only in accordance with thefollowing claims appended hereto and their equivalents.

1. A mobile platform comprising: a flight control system generatingcommands; a structure; an aerodynamic surface; an actuator operablycoupled to the surface to control the surface, a vibration sensorcoupled to the structure and generating a signal representative of avibration of the structure; and a circuit coupled to the actuator andincluding, a first input communicating with the flight control system toaccept the commands; a second input communicating with the vibrationsensor to accept the vibration signal, and a summing elementcommunicating with the inputs to sum the signal and the command, thecircuit controlling the actuator with the summed signal and command. 2.The mobile platform according to claim 1, further comprising a positionsensor generating a signal representative of the position of theactuator, a filter in communication with the position sensor and thevibration sensor and filtering the vibration signal from the positionsignal, the filter including an output and communicating the filteredposition signal to the flight control system via the output.
 3. Themobile platform according to claim 1, further comprising a first powercable connected to the actuator and supplying power to the actuator, theactuator including a second power cable communicating with the circuitand the first power cable.
 4. The mobile platform according to claim 1,wherein the mobile platform is an aircraft.
 5. The mobile platformaccording to claim 4, wherein the structure is a portion of a wing. 6.The mobile platform according to claim 4, wherein the structure is ahousing of the actuator.
 7. The mobile platform according to claim 4,wherein the surface is an aileron.
 8. The mobile platform according toclaim 7, wherein the actuator is at an outboard location of the aileron.9. A mobile platform comprising: a control system generating commands; astructure; an aerodynamic surface; an actuator operably coupled to theaerodynamic surface to control the aerodynamic surface; a vibrationsensor coupled to the structure and generating a vibration signalrepresentative of a vibration of the structure; and a circuit coupled tothe actuator adapted to communicate with the control system and thevibration sensor and to use the commands and the vibration signal tocontrol the actuator.
 10. The mobile platform of claim 9, wherein themobile platform comprises an airborne mobile platform.
 11. The mobileplatform of claim 9, wherein the mobile platform comprises an aircraft.12. The mobile platform of claim 9, wherein the circuit comprises: afirst input communicating with the control system to accept thecommands; a second input communicating with the vibration sensor toaccept the vibration signal; and a summing element communicating withthe inputs to sum the signal and the command.
 13. The mobile platform ofclaim 12, wherein the circuit comprises a third input for accepting aposition signal representative of a position of the actuator.
 14. Themobile platform of claim 13, wherein the circuit further comprises afilter in communication with the second and third inputs adapted tofilter the vibration signal from the third signal.
 15. The mobileplatform of claim 9, further comprising a first power cable connected tothe actuator and supplying power to the actuator.
 16. The mobileplatform of claim, 15, wherein the actuator includes a second powercable communicating with the circuit and the first power cable.
 17. Themobile platform of claim 9, wherein the structure comprises a wing. 18.An airborne mobile platform comprising: a flight control systemgenerating commands; a structure that forms a portion of the airbornemobile platform; an aerodynamic surface; an actuator operably coupled tothe aerodynamic surface to control the aerodynamic surface; a vibrationsensor coupled to the structure and generating a vibration signalrepresentative of a vibration of the structure; and a circuit coupled tothe actuator adapted to communicate with the control system and thevibration sensor and to use the commands and the vibration signal tocontrol the actuator; and the circuit including: a position sensor incommunication with the vibration sensor that senses a position of theactuator; a filter in communication with the position sensor and thevibration sensor, and adapted to filter the vibration signal from theposition signal, and to output a filtered position signal to controlsaid flight control system.
 19. The airborne mobile platform of claim18, wherein the structure is a portion of a wing.
 20. The airbornemobile platform of claim 18, wherein the surface is an aileron.